The present invention relates to a turbine engine, and more particularly to a tip turbine engine with multiple fan stages and/or multiple tip turbine stages.
An aircraft gas turbine engine of the conventional turbofan type generally includes a forward bypass fan, a low pressure compressor, a high pressure compressor, a combustor, a high pressure turbine, and a low pressure turbine, all located along a common longitudinal axis. The low and high pressure compressors are rotatably driven to compress entering air to a relatively high pressure. This high pressure air is then mixed with fuel in the combustor, where it is ignited to form a high energy gas stream. This gas stream flows axially aft to rotatably drive the high pressure turbine, which rotatably drives the high pressure compressor via a high spool shaft. The gas stream leaving the high pressure turbine is expanded through the low pressure turbine, which rotatably drives the forward bypass fan and low pressure compressor via a low spool shaft.
Although highly efficient, conventional turbofan engines operate in an axial flow relationship. The axial flow relationship results in a relatively complicated elongated engine structure of considerable longitudinal length relative to the engine diameter. This elongated shape may complicate or prevent packaging of the engine into particular applications.
A recent development in gas turbine engines is the tip turbine engine. Tip turbine engines include hollow fan blades that receive core airflow therethrough such that the hollow fan blades operate as centrifugal compressors. Compressed core airflow from the hollow fan blades is mixed with fuel in an annular combustor, where it is ignited to form a high energy gas stream which drives the turbine that is integrated onto the tips of the hollow bypass fan blades for rotation therewith as generally disclosed in U.S. Patent Application Publication Nos.: 20030192303; 20030192304; and 20040025490. The tip turbine engine provides a thrust-to-weight ratio equivalent to or greater than conventional turbofan engines of the same class, but within a package of significantly shorter length.
In some tip turbine engine designs, core airflow may be compressed by an axial compressor before entering the hollow fan blades for further, centrifugal compression. The axial compressor may include an axial compressor rotor with one or more stages of radially-extending compressor blades alternated with stages of static compressor vanes. Increasing the number of stages of compressor blades and compressor vanes increases the compression of the core airflow and the efficiency of the engine, but increases the overall length and weight of the engine and the number of parts.